The present invention relates to abradable surfaces for high temperature applications, and more particularly relates to such surfaces on ceramic matrix composite (CMC) ring segments for combustion turbines.
Some components of combustion turbines operate at high temperatures, and thus may require thermal barrier coatings (TBCs). Conventional TBCs typically comprise a thin layer of zirconia or other ceramic material. In some applications, the coatings must be erosion resistant and also abradable. Turbine ring seal segments must withstand erosion and must also have tight tolerances on a radially inward sealing surface opposed the tips of rotating turbine blades. To minimize these tolerances, the sealing surface of ring segments may be made abradable in order to reduce damage to the turbine blades upon occasional brushing contact of blade tips with the sealing surface.
Improvements in gas turbine efficiency rely on breakthroughs in several key technologies as well as enhancements to a broad range of current technologies. One of the key issues is a need to tightly control rotating blade tip clearance. This requires that turbine ring segments are able to absorb mechanical rubbing by rotating blade tips.
For modern conventionally cooled and closed loop steam cooled turbine ring segments, a thick thermal barrier coating of about 0.1 inch on the ring segment surface is required for rubbing purposes. The latest advanced gas turbine has a hot spot gas temperature of over 1,500 degrees C. at the first stage ring segment. Under such heat, a TBC surface temperature of over 1,300 degrees C. is expected. Thus, a conventional abradable TBC is no longer applicable because conventional TBCs are typically limited to a maximum surface temperature of about 1,150 degrees C.
Friable graded insulation (FGI) materials are disclosed in U.S. Pat. No. 6,641,907 commonly owned by the present assignee. The effectiveness of FGI as an abradable refractory coating is based upon control of macroscopic porosity in the FGI to deliver acceptable abradability. Such a coating may consist of hollow ceramic spheres in a matrix of alumina or aluminum phosphate. To bond an FGI layer to a metal ring segment, an FGI-filled metallic honeycomb structure has been proposed in U.S. Pat. No. 6,846,574 commonly owned by the present assignee. In this technique a high temperature metal alloy honeycomb is brazed to the metallic substrate. The honeycomb, once oxidized prior to FGI application, serves as a mechanical anchor and compliant bond surface for an FGI filler, and provides increased surface area for bonding.
Further advances in high temperature abradable surfaces for gas turbine ring segment surfaces are desired.
The invention is explained in the following description in view of the drawings that show:
A gas turbine component, especially a ceramic matrix composite (CMC) ring segment, is described herein with an abradable surface exposed to a hot gas flow. In contrast to prior art, no thermal barrier coating is applied to the exposed surface. Instead, the CMC itself is used as its own thermal barrier, but is modified to allow for abradability. The current invention provides an array of depressions directly in the CMC surface to increase its abradability, allowing occasional brushing contact with turbine blade tips with reduced wear on the blade tips. This technology is especially applicable to CMC ring segment walls formed by laminate construction, in which CMC layers are oriented edgewise in a stacked configuration.
Behavior of CMC exposed to high temperatures shows reduction in strength over long periods; however such a reduction in strength should not be limiting for the present invention because strength is not the material property of primary concern for a wear surface. Since a CMC surface 34, 44 in this invention is directly exposed to the hot working gas, it will be exposed to temperatures over 1200° C. This will reduce its strength but will also increase its hardness. The increase in hardness will beneficially reduce erosion of the surface. The surface may be allowed to age during operation of the gas turbine engine, or it may be pre-aged prior to being placed into operation. A thin, hard ceramic coating, for example alumina, may be applied to the CMC edges as temporary erosion protection until CMC hardening occurs.
The present invention eliminates the need for an abradable thermal barrier coating such as FGI, thus eliminating the associated bond joint and avoiding any concern about differential elasticity between the two materials. Accordingly, the invention is expected to provide improved component reliability and durability and reduced manufacturing expense compared to prior art coating methods.
While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.
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Number | Date | Country | |
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20080279678 A1 | Nov 2008 | US |